1. Field of the Invention
The present invention generally relates to integrated solid state electrical power sources and to structures and vehicles containing such integrated power sources embedded in or formed as a structural member. More particularly, the invention relates to solar power systems containing flexible integrated photovoltaic solar cells and advanced energy storage systems in which ultra thin-film batteries, capacitors and other components are formed as a single unit and embedded into a supporting or covering element of a structure.
2. Description of Related Art
Integrated Power Sources. Several integrated electrical power sources have been devised over the past two decades for a variety of applications, and in some cases a battery and/or a solar photovoltaic collection device has been integrated or embedded into a support or covering for a structure or device. The term “integrated power source” refers to a power source that is structurally combined as a single unit. For example, U.S. Pat. No. 4,740,431 describes an integrated solar cell and battery primarily for use in portable electronic devices such as radio transceivers, portable computers and emergency lights, as integrated power generation and storage modules. Thin film lithium-aluminum alloy anodes, polyethylene oxide/lithium salt solid electrolyte, and a molybdenum selenide cathode are described as exemplary battery components and the solar cells may contain amorphous silicon.
U.S. Pat. No. 5,626,976 describes a certain flexible energy storage device with integral charging unit (photovoltaic device). Conventional aqueous or non-aqueous electrolyte or polymer gel electrolyte material, or a solid state electrolyte is employed, and enclosed in a polymeric vapor barrier package.
U.S. Pat. No. 5,180,645 describes a certain integral solid state embedded power supply for a self-powered portable electronic product, such as a radio. The elimination of the outside metallic case of each cell and the elimination of the outer packaging for the overall battery provides a considerable reduction in volume and weight of the product.
U.S. Pat. No. 5,644,207 describes a renewable modular integrated power source that is bonded to a housing or structure or is molded into a desired shape using the battery material itself. In one application solar cells backed with a thin film polymer battery supply a lightweight fully integrated power source. Conventional solid state or ionically conducting gel polymer electrolyte batteries, such as that described in U.S. Pat. No. 5,637,421 are employed.
U.S. Pat. No. 5,695,885 describes a battery configured in the form of a flexible wrist band for a watch or personal radio, television or communication device, with a plurality of photovoltaic cells disposed on the outer surface of the battery. Conventional alkaline or Ni—Cd type batteries are described along with presently known flexible anode and cathode materials.
U.S. Pat. No. 6,224,016 describes a flexible energy producing envelope material or covering for a balloon used in high altitude and stratospheric applications. The covering includes a flexible solar cell layer, a flexible substrate that matches the shape and size of the airship gore as well and an electrically conductive conduit disposed in a flexible electrically non-conductive adhesive connecting the flexible solar cell substrate to the airship substrate. No energy storage means is integrated into the covering with the photovoltaic cell layer, however.
For powering satellites, aerobots (balloons) and other spacecraft applications, such as altitude control systems, communications and various payloads, a conventional integrated power system (e.g., solar power system) that might be satisfactory for powering a small, personal device such as a radio or a wristwatch would not be adequate, particularly spacecraft intended for long missions in space.
Space Power Systems. A number of advanced propulsion and energy storage technologies with reduced mass and volume, long service life, higher reliability, thermal and radiation resistant and low cost for use in spacecraft such as satellites, nanosats and aerobots (balloons) are required for future space missions, and for military, surveillance, scientific research and commercial applications. For instance, there is a great need for a longer lasting power source for use in deep space exploration away from the sun, and no suitable power source is presently available using today's technology. In the cryogenic conditions of deep space, aerobots or balloons will require more battery capacity and higher power capabilities in a lightweight design than is presently possible with existing power sources. Typically, nanosats and aerobots call for the incorporation of photovoltaic devices on the surface of panels to charge the auxiliary on-board batteries for a variety of usages. Another application for lightweight sources of electrical energy is in the construction of lightweight high altitude or spaceborne platforms (e.g., for missile defense use).
Further complicating the problem, today's satellite designers continually strive to raise the level of onboard power generation while at the same time endeavoring to lower the cost of such increased power capability. Therefore, satellites that are now envisioned for various future anti-missile roles, for example, have an ever increasing demand for power. Although the goals and operational requirements of satellite missions can vary widely, they are in almost every case constrained by the type, level, and duration of the on-board power source.
Space power systems designed for continuous high-power-output applications have most commonly been based on the use of photovoltaic panels that regenerate the battery. In all orbits, satellites are subjected to a greater or lesser number of eclipses, and thus have to rely on internal energy to continue their missions. This internal energy is provided by rechargeable batteries, which store the surplus energy generated by spacecraft solar arrays. In geosynchronous orbit (GEO), the satellite is in eclipse for only a short time: for approximately 4 months, the satellite is in the sun at all times; for the next 50 days, the satellite is in eclipse for periods up to 72 minutes per day (it builds from 0 minutes to 72 minutes in the first 25 days and then drops down to 0 minutes); then, the satellite is in the sun for approximately 4 months. GEO applications require approximately 100 charge/discharge cycles (C) per year and as such the requirement is not so strenuous. In low earth orbit (LEO) missions, the battery typically provides power for 30 minutes at the 1 C rate, followed by a one hour charge. LEO applications usually require about 6000 cycles per year.
Ni—H2 Batteries. In the past decade, nickel-hydrogen batteries have been the technology of choice for both commercial and defense-related satellites in both GEO and LEO applications. They have inherent advantages over their predecessor nickel-cadmium batteries. These include superior energy density, longer cycle life, and better tolerance to overcharge and reversal. The goal of increasing cell energy density provided the original impetus for the development of the nickel-hydrogen system. In a battery, the useable energy density of a nickel-cadmium battery is typically in the range of 20–35 Wh/kg while that of a nickel-hydrogen battery it is about 40–45 Wh/kg.
Present day satellites can be categorized into four groups: large, small, mini and micro. The large satellites have mass greater than 1000 kg; the small category is 500–1000 kg; the mini in the 100–500 kg range and the micro less than 100 kg. Table 1 shows satellite battery-power systems data for some previous space missions. The power requirements for GEO satellites are in the 10–15 kW range for extremely long-term missions lasting up to 15 years.
TABLE 1Specifications of Some Previous Satellite Battery-Power SystemsArrayBatterySatellitePowerBatterySizeSatelliteProgramPayloadMass (kg)(W)Type(Ah)HESSINASA SmallSolar Imager290505Ni—H215ExplorerCPVMightySatAFRLResearch &121330NiCd102.1ImagingDeep Space 1JPL NewAdv. Tech. Ion4862500Ni—H230MilleniumProp.CPVARGOSAFExperiments21001074Ni—H245IPVStardustNASA DiscoverySample returns380Ni—H216CPVNEARNASA DiscoveryAsteroid8051800Super9RendevousNiCdEO-1GSFC NewEarth529600Super50MilleniumobservationsNiCdTerriersNASA STEDIEarth atm.12531NiCd4.8observationsHETE-2NASAGamma ray124168NiCd1.2detection.
For longer missions requiring high cycle life and cryogenic operating conditions, low energy density Ni—H2 batteries are now employed. These batteries are not only bulky and constructed in heavy steel pressure vessels, but also offer lower current capabilities. For shorter missions and low cycle life requirements in which high power is required, NiCd batteries are typically used. These batteries and solar panels occupy between 15–25% of the total weight and volume of a present day satellite. The volume of the satellite is a function of the surface area for the solar panels and hence the power availability. They vary from about 90 m3 for satellites weighing over 2000 kg to about 2.5 m3 for those weighing about 35 kg.
Lithium Ion Batteries. Space power systems of the future are projected to require power levels that may extend far beyond the current levels of demand. Thus, there is an increasing need for lightweight, high energy density batteries with long active and cycle lives beyond what can be delivered by nickel-hydrogen batteries. Toward that end, lithium ion battery systems are at this time undergoing intense investigation. For Example, U.S. Pat. No. 5,456,000 describes one type of lithium ion rechargeable battery containing polymeric film composition electrodes and separator membranes. An advantage of a lithium ion battery system is that the useable specific energy is two to three times greater than that of nickel-hydrogen batteries. This represents a significant launch cost savings or increased payload mass capabilities. However, the energy density and cycle life of a conventional lithium ion battery system is typically only about 150 watt hours per kilogram and 500 cycles, respectively, under deep discharge conditions.
Solid Polymer Electrolyte Batteries. Commonly assigned U.S. Pat. Nos. 6,645,675 and 6,664,006, and PCT International Publication Nos. WO 01/17051 and WO 01/17052 describe all-solid-state electrochemical cells and batteries with a flexible, ionically conductive polymer membrane as the electrolyte For example, a rechargeable all-solid-state lithium polymer electrolyte battery comprises an ultra thin lithium anode, which may be either a metallic lithium element or a lithium metal layer about 0.1 to 100 microns thick, over the metallization layer of a metallized polymer substrate. The metallized polymer substrate has an inactive polymer layer about 0.5 to 50 microns thick and a metallization layer 0.1–1μ thick on top of the inactive polymer layer. This battery also has an ultra thin-film cathode layer containing a metallized polymer substrate. The metallized polymer substrate has an inactive polymer layer about 0.5 to 50 microns thick and a metallization layer about 0.01–1μ thick on top of the inactive polymer layer, and has a layer of active cathode material 0.1–100μ thick on top of the metallization layer. The battery also has a polymer electrolyte layer 0.2–100μ thick placed between the above-described anode and cathode layers. This polymer electrolyte has a conductivity greater than 1×10−4 S/cm at 25° C., or may even conduct as well below 25° C. The polymer electrolyte comprises a mixture of a base polymer material comprising at least one ionically conductive polymer and having an initial conductivity of at least about 1×10−5 S/cm at 25° C. when combined with a lithium salt. The mixture also includes the lithium salt, an inorganic filler having an average particle size <0.05 micron in diameter and a surface area of at least about 100 m2/g, and a lithium ion conducting material having an average particle size <0.1 micron in diameter and an initial ionic conductivity of at least 2×10−3 S/cm at 25° C. In some embodiments, the inorganic filler is 0.1–20% (by volume of solid polymer electrolyte) high surface area filler having an average particle size ≦0.01 micron and chosen from the group consisting of fumed silica and alumina. In some embodiments, the lithium ion conductor material is 0.1–80% sulfide glass (by volume of solid polymer electrolyte). In some embodiments, the lithium ion conductor material is a ceramic ion conductor chosen from the group consisting of lithium beta alumina and silicates. Ion mobility is achieved through coordination of electrolyte ions by sites on the polymer chain, thus promoting electrolyte dissolution and salt dissociation. Such a battery design overcomes the disadvantages inherent in liquid electrolytes and provides better long-term storage stability. By also employing thin, flexible electrodes, such batteries can be made into virtually any shape and size, are reasonably rugged and leakproof, have high specific energy (Wh/kg) (gravimetric) and energy density (Wh/L) (volumetric), high cycle life, low self-discharge, high current drain capability, lower resistance, and wider operating temperature range.
Two major factors that drive satellite design are launch costs per kilogram of satellite and instruments and power availability. Since small and moderate mission costs can run from $100 million to greater than $1 billion, reliability, efficiency, and density of power in the satellite design and components is necessary. Moreover, future satellites for missile defense applications will require not only greater longevity from the power source but also higher power, lower weight and volume, greater degree of power management, and significantly lower cost, so that more firepower can be packed into these satellites. Other applications such as surveillance, deep space exploration or terrestrial use will have similar requirements. The existing space power technologies are only capable of providing power sources that are rigid, bulky, heavy and costly, and which are monofunctional, i.e., they serve as a power source only. Present day satellites also need, in addition to the power source components, a special ‘skin’ enclosure that is sturdy enough to protect the power source from the sun's heat and from space debris. This is necessary since the conventional batteries contain a liquid electrolyte that can evaporate away at high heat or cause a dangerous situation. At the same time, the electrodes need to be contained in a rigid environment to prevent them from disintegrating due to constant bombardment by space debris. Furthermore, the conventional liquid electrolyte batteries, and even the newer lithium ion batteries, do not provide satisfactory high current drain capabilities for use in many of the latest applications being developed, such as anti-missile systems, aerobots or balloons for planetary exploration.
Another problem with existing photovoltaic systems is that the solar cells at optimum levels of performance get hot, which reduces the efficiency. Solar cells operate better at cooler temperatures.
Despite the technological advancements provided by prior art devices, known integrated power sources still suffer from limitations of excessive size and weight, lack of sufficient flexibility or conformability, insufficient power density (watts/volume) and specific power (watts/weight), particularly for use in the new and future high altitude and space-related military programs, for space exploration, and for various other scientific research and commercial applications.